1. Field of the Invention
The present invention relates generally to a gas turbine engine, and more specifically to an air cooled turbine airfoil with leading edge film cooling.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
In a gas turbine engine, such as a large frame heavy-duty industrial gas turbine (IGT) engine, a hot gas stream generated in a combustor is passed through a turbine to produce mechanical work. The turbine includes one or more rows or stages of stator vanes and rotor blades that react with the hot gas stream in a progressively decreasing temperature. The efficiency of the turbine—and therefore the engine—can be increased by passing a higher temperature gas stream into the turbine. However, the turbine inlet temperature is limited to the material properties of the turbine, especially the first stage vanes and blades, and an amount of cooling capability for these first stage airfoils.
The first stage rotor blade and stator vanes are exposed to the highest gas stream temperatures, with the temperature gradually decreasing as the gas stream passes through the turbine stages. The first and second stage airfoils (blades and vanes) must be cooled by passing cooling air through internal cooling passages and discharging the cooling air through film cooling holes to provide a blanket layer of cooling air to protect the hot metal surface from the hot gas stream.
The leading edge of the airfoil is exposed to the highest gas flow temperature and therefore requires the most amount of cooling. FIGS. 1 and 2 shows a prior art turbine blade with the leading edge region being cooled using a showerhead arrangement of film cooling holes 11 and two rows of gill holes (14). Cooling air is delivered to a supply channel and flows through a row of metering and impingement holes 13 to produce impingement cooling on the backside surface of the leading edge wall. The spent impingement cooling air then flows through the film cooling holes and gill holes to provide a layer of film cooling air on the outer surface of the leading edge region.
The showerhead film cooling holes 11 are supplied with cooling air from a common impingement channel (12) and discharged at various gas side pressures. Because of this prior art design, the cooling flow distribution and pressure ratio across the showerhead film holes for the pressure and suction side film rows is predetermined by the impingement channel pressure. Also, the standard film holes pass straight through the airfoil wall at a constant diameter and exit the airfoil at an angle to the surface. Some of the coolant is subsequently injected directly into the mainstream gas flow causing turbulence, coolant dilution and loss of downstream film cooling effectiveness. And, the film hole breakout on the airfoil surface may induce stress issues in the blade cooling application.
The prior art blade includes three rows of film holes in the showerhead arrangement. The middle row of film holes is positioned at the airfoil stagnation point where the highest heat loads is located on the airfoil leading edge region. Film cooling holes for each film row are inclined at 20 to 35 degrees toward the blade tip as seen in FIG. 3. A major disadvantage of this prior art design is an over-lapping of film cooling air ejection flow in a rotational environment (in a rotor blade) and low through-wall convection area as well as heat transfer augmentation. This prior art film cooling hole arrangement and design results in the appearance of hot streaks 16 on the airfoil surface because the cooling air flow from the middle row flows over the film cooling holes on the outer rows without flow over the space between adjacent film holes in the same row as seen in FIG. 4.
The prior art blade with showerhead film cooling holes is formed by an investment casting process that uses a ceramic core to form the internal cooling air passages and features. The film cooling holes are then drilled into the solid metal blade using a process such as laser drilling or EDM drilling. Because of the limitations of the ceramic core is forming cooling air passages and features, hole diameters are limited to no smaller than around 1.3 mm because the ceramic piece would break when the liquid metal flows around the ceramic core.